Apparatus and method for determining information for aircraft

ABSTRACT

The present invention provides a method for determining information for an air vehicle including the steps of providing a first coefficient of pressure data point, providing a situation involving the air vehicle, deriving pressures during the situation and deriving a second coefficient of pressure data point from the pressures derived during the situation, and defining a line passing through the first and second coefficient of pressure data points, whereby a selected coefficient of pressure data point on the line corresponds to the information. The invention encompasses apparatus for accomplishing the method. In some embodiments, the information relates to angles from zero lift and/or zero yaw.

[0001] The present application claims the priority of a U.S. provisionalpatent application, Ser. No. 60/067,147, filed Dec. 2, 1997, and a U.S.non-provisional patent application, Ser. No. 09/201,067, filed Nov. 30,1998.

BACKGROUND OF THE INVENTION

[0002] The present invention relates to flight instrumentation foraircraft. More particularly, it relates to a measurement, datamanipulation and information display system and method for use on airvehicles, wherein the system includes pressure sensors, a dataprocessor, at least one data port for electrically exporting angles fromzero lift, and an indicator for communicating information relating toangles from zero lift.

[0003] Past and present pressure type angle of attack devices use avariety of differently shaped probes. A hemispherical end sensor isdescribed in U.S. Pat. No. 3,318,146. Angle of attack (AOA) and angle ofside slip both are calculated from the pressure signals present at fiveports on the end. This angle of attack measuring device requires use ofthe probe as described and the probe must be installed at a very preciseangle relative to the cord of the wing or some other longitudinal planeof the air machine.

[0004] Sharp tipped pitot static tubes with angle of attack sensingports are shown in U.S. Pat. Nos. 4,096,744 and 4,378,696. These patentsshow probes providing differential pressures which are used fordetermining angle of attack. In the '696 patent, a combination ofdifferential pressure at two ports on the probe, measured pitot andstatic pressure is used. In the '744 patent, measured pitot pressure,pressure at one of the angle of attack sensitive ports and measuredimpact pressure are utilized for the calculations.

[0005] U.S. Pat. No. 5,616,861 discloses a plurality of air sensingprobes symmetrically mounted on opposite sides of a vertical centerplane of the air vehicle which includes the longitudinal axis of the airvehicle.

[0006] In each of the above instances, the angle of attack systemsrequire special probes usually installed at very precise angles relativeto the cord of the airfoil, each other, or other longitudinal axis ofthe air vehicle.

[0007] Additionally, in the above instances and other known sensorsystems, the probes must be located well ahead of the wing to reduce theeffect of changes in lift and configuration. Usually the probe islocated on or near the nose of an aircraft. This is a problem since manyfactors can affect the relationship between the local AOA and the truewing AOA. The airflow around the nose is not the same as that at thewing. Pitch rate errors occur when the nose of an aircraft is pitched upor down causing the probe to indicate too high or too low. In a turn,for example, the nose is pitched up reducing the local flow angle andcausing the AOA reading to be too low. Because of the shape of the nose,sensitivity errors are introduced. A one degree change in true wing AOAmay cause the local flow angle at the nose to be 1.5 to 2 degrees.

[0008] Another problem with systems such as those noted above is thatthe measurement used was based on angle of attack which is the anglebetween the cord of the airfoil and the relative wind. The abovepatented devices are representative of the many devices that have beeninvented and installed on aircraft to produce a known pressuredifferential at a known angle of attack. Typically, they require that aprobe or probes be installed at a fixed angle relative to a known planeof the air vehicle, usually the cord or surface of the wing orlongitudinal axis of the air vehicle.

[0009] In recent years, there has been emphasis on making pressuresensing probes compact, light and with low drag as illustrated by thedevices shown in U.S. Pat. Nos. 4,836,019, 5,616,861 and 5,331,849.However, there is still room for improvement.

[0010] U.S. Pat. No. 4,350,314 describes a stall condition detector thatuses four pressure ports in the wing which are ported to a capacitancedevice within the wing. The detector involves four electricallyconductive hollow tubes with rods extending up through the tubes. Thetubes are filled with a dielectric fluid. Devices of this type may besubject to errors induced by temperature and other forces acting uponthe dielectric fluids resulting from turbulence, G loads, slips andskids. Additionally, with the emphasis of higher loaded airfoils, thenewer airfoil designs are smaller and thinner, thus having smallerinterior compartments, making these devices difficult to install. Inaddition, the angle of attack display of this type of device does nottake into account high lift devices which may significantly reconfigurethe shape of the airfoil.

[0011] An aircraft instrument system for communicating accurateinformation involving angles from zero lift, wherein the information maybe derived from simple pressure ports rather than known probes ordetectors of the type disclosed in the '314 patent, would beadvantageous.

SUMMARY OF THE INVENTION

[0012] The present invention relates to a measurement and display systemfor use on air vehicles having an indicator visibly displaying anglesfrom zero lift and having a data port for electrically exporting anglesfrom zero lift. In one embodiment, pressures from upper and lower wingports and pressures from the typical pitot and static ports of an airvehicle are converted to digital data and mathematically divided oneinto the other. The result is a coefficient of pressure (CP) whichvaries uniquely with angles from zero lift of the air vehicle. Theangular deviation from angle of zero lift display is designed for airvehicles and is located in the cockpit and displays angles from zerolift on a digital display. The data port electrically exports anglesfrom zero lift to other electronic devices using a compatiblecommunications port. It has been discovered that if the coefficient ofpressure data point for zero degrees angular deviation from angle ofzero lift and a minimum of one other coefficient of pressure data pointis stored permanently in the system's non-volatile memory, other anglesfrom zero lift can be determined given any coefficient of pressure. Forany given airfoil it has been discovered that a minimum of two or morecoefficient of pressures will accurately define the relationship betweenthe coefficient of pressure and the angular deviation from angle of zerolift. Once the data points are stored in the system's non-volatilememory, all other coefficient of pressures can be equated to a specificangular deviation from angle of zero lift. Thus, in one embodiment, thepresent invention provides an apparatus and method for measuringpressures, for deriving and storing in non-volatile memory two or morecoefficient of pressure data points, and for exporting and displayingthe angles from zero lift.

[0013] An advantage of the present invention is that a microprocessormay be used to permanently store the two or more coefficient ofpressures which may be copied and installed in the angular deviationfrom angle of zero lift system of similar air vehicles of the same typeto accurately display angles from zero lift.

[0014] The instrument of the present invention could also measureangle-of attack (AOA) which is the angle between the relative wind andthe chord of the airfoil. But the zero reference for AOA is of no valueto the pilot and has no aerodynamic significance. In the case ofmeasuring angles from zero lift, the reference zero angle from zero liftis the angle where the air vehicle creates no lift and therefor is theonly angle where the induced drag is zero. For symmetrical airfoils, AOAand angles from zero lift are the same. For a nonsymmetrical airfoil,AOA and angles from zero lift vary by a fixed constant.

[0015] The present invention provides an apparatus and method for use onaircraft and includes pressure sensors, a processor, at least one dataport for electrically exporting angles from zero lift, and an indicatorfor visibly displaying angles from zero lift. The present inventionprovides the apparatus and method for measuring angular deviation fromangle of zero lift using pressures from pressure sensing ports,assembled with pressure sensors that minimize the need for temperaturecompensation and a processor that computes angular deviation from angleof zero lift using two or more flight data points permanently recordedinto non-volatile memory for a variety of airfoil configurations. In oneembodiment, upper and lower airfoil pressure ports ported to an angulardeviation from angle of zero lift processor may be simply a small holeor holes bussed together in the upper surfaces of the wing and/or asmall hole or holes bussed together in the lower surfaces of the wingand connected to tubes routed to the processor. In this embodiment,there would be no need for probes such as those described in theabove-noted patents. For example, the pressure ports for the pitot andstatic pressures may be from a pitot/static tube or individual pitot andstatic ports of the type described in the above-noted patents, or may befrom the air vehicles airspeed pressure tubes, eliminating the need fora separate pitot static probe.

[0016] It will be noted that the existence and use of pressure sensorsintegrated on silicone, which are generally trouble free, make thedevice according to the present invention particularly advantageous. Inthe case where known type probes are too heavy or expensive, it isnevertheless possible to use a pressure port in the top of the airfoilwith a water separator and, optionally, a pressure port in the bottom ofthe airfoil to provide upper and lower wing local pressures.

[0017] In the case of using upper and lower local wing pressures, in oneembodiment, the wing pressures can be ported to a single differentialpressure transducer. The pitot and static pressures can be ported to asingle differential pressure transducer. No matter where the source ofthe wing local differential pressure is, the result of dividing the winglocal differential pressure by the pitot/static differential pressure(dynamic pressure) provides a coefficient of pressure (CP) which variesuniquely with angles from zero lift for the air vehicle.

[0018] It should be appreciated that, in the case of using either aupper or lower pressure port or ports bussed together, the wingport/ports can be ported to a single absolute pressure sensor, the pitotpressure can be ported to a single absolute pressure transducer, and thestatic pressure can be ported to a single absolute pressure transducer.The result of computing the difference between the wing pressure and thestatic pressure and dividing by the difference between the pitotpressure and the static pressure is a coefficient of pressure whichvaries uniquely with angles from zero lift for the air vehicle.

[0019] The angle from zero lift processor converts the three or fourported pressures to an analog output using temperature compensated solidstate pressure transducers. An analog to digital converter may be usedto convert the analog data to digital data. The digital pressure datamay be processed by a microprocessor which computes a coefficient ofpressure. The coefficient of pressure may be used to calculate theangular deviation from angle of zero lift using a known relationshipbetween the coefficient of pressure and the angular deviation from angleof zero lift.

[0020] The relationship between the coefficient of pressure and theangle from zero lift is determined by using two or more data pointssaved in the non-volatile memory of the angle from zero lift computer.These permanently saved calibration points may be derived by flying theair vehicle into two or more flight maneuvers. A zero G maneuver thatdetermines the coefficient of pressure where the angular deviation fromangle of zero lift is zero degrees is saved to non-volatile memory. Ahigher angle or angles from zero lift are flown and saved tonon-volatile memory. In the case of two data points, these pointsmathematically determine points along an assumed relationship (linearand/or curvilinear) between the coefficient of pressure and angles fromzero lift. In the case of more than two data points, these pointsmathematically determine points along the assumed smoothed linerelationship between the coefficient of pressure and angles from zerolift.

[0021] The angles from zero lift are displayed on a angle from zero liftdisplay instrument and angles from zero lift are exported via a dataport generally compatible with flight data computers. When the angulardeviation from angle of zero lift approaches a critical angle, such asthe angle at which airflow will separate from the airfoil, the systemexports electrical, aural and/or visual warnings via the data port andthe visual display or audio system.

[0022] An advantage of the present invention is that it does not requirethe precise installation of any device or probe that produces a knownpressure at a known angle of attack thereby eliminating all installationcalibrating efforts and errors. The present invention uses a zero Gflight maneuver to determine when the airfoil is producing no lift andrecords the coefficient of pressure in non-volatile memory of a suitablecomputer or microprocessor. The zero G maneuver determines thecoefficient of pressure for the zero degree angle from zero lift. Oneother data point or more at any higher angle from zero lift is obtained,and the relationship between the coefficient of pressure and angle fromzero lift is defined for a specific airfoil configuration. The anglefrom zero lift system instrument's processor uses various acceptablemathematical functions to connect the points recorded in non-volatilememory providing the angular deviation from angle of zero lift for anycoefficient of pressure resulting from any combination of pressures atthe pressure ports for a known airfoil configuration.

[0023] An object of the invention is to provide an angular deviationfrom angle of zero lift measurement system for air vehicles using sensedpressures, and a display of angles from zero lift information on adisplay in the cockpit of the air vehicle. Another object of theinvention is to eliminate the requirement for specialized probes andbulky sensors thus reducing weight, drag, complexity and cost. Anotherobject of the invention is to provide a system that will produce angulardeviation from angle of zero lift data on any air machine without theinstallation and calibration requirements for aligning a probe or portrelative to a plane of the air vehicle. Another object of the inventionis to provide angular deviation from angle of zero lift information tothe air vehicle pilot and warnings when the angular deviation from angleof zero lift is at critical angles such as warnings when the angle fromzero lift is approaching an unsafe angle from zero lift where theairflow will separate from the airfoil. Another object of the inventionis to provide a visual indication of when the airfoil is operating atthe optimum angular deviation from angle of zero lift for an approach,or for the best lift over drag or other optimum performance parametersbased on AOA. Another object of the invention is to warn a pilot whenthe air vehicle's airfoil is not properly configured for an approach tolanding and the pilot intends to land.

[0024] Another object of the present invention is to eliminate the needto de-ice or anti-ice the probe or vane used on other AOA devices. Stillanother object is to eliminate or reduce the radar profile and/or theheat profile. These objects are accomplished by the present inventiondue to the nonobtrusive ports and their flush location relative to the awing surface which requires no de-icing or anti-ice treatment.

[0025] An advantage of the present invention is that it may be used tocreate and develop coefficient of pressure data points based on a windtunnel or other simulation situation, or by using CFD (computationalairfoil data analysis). Additional advantages of the present inventionare that the apparatus is light in weight and easy to install, it doesnot require the use of ugly, protruding probes, it does not requireprecise alignment of probes, it does not require moving parts, it can beadapted to virtually any aircraft, including drones, remote pilotedvehicles (RPV's) and the like, and its operation and use is easy tounderstand and learn.

[0026] Other features and advantages of the apparatus and method of thepresent invention will become more fully apparent and understood withreference to the following description and drawings, and the appendedclaims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0027]FIG. 1 is a fragmentary elevational view of a portion of an airvehicle depicting two pressure detecting orifices located on the airfoilthereof.

[0028]FIG. 2 is an enlarged central cross-sectional view depicting oneembodiment of a pressure port attached to an airfoil lower interiorsurface.

[0029]FIG. 3 is an enlarged central cross-sectional view depicting oneembodiment of a pressure port and air water separator attached to anairfoil upper interior surface.

[0030]FIG. 4 is block diagram depicting one embodiment of the processused to collect the air pressures, compute the coefficient of pressures,store the coefficient of pressures in non-volatile memory, and outputangles from zero lift.

[0031]FIG. 5 is a block diagram depicting electronic instrumentationused in one embodiment of the invention to compute the coefficient ofpressure and the angular deviation from angle of zero lift.

[0032]FIGS. 6a-d depicts a front view of one embodiment of an angle fromzero lift indicator which visibly displays the angular deviation fromangle of zero lift in an analog bar display and a digital display of theactual angle in units.

[0033]FIG. 7 is a substantially representational view depicting theinstallation of one embodiment of the present invention.

[0034]FIG. 8 is a substantially representational view depicting oneembodiment of components of the present invention, including thedisplay, outputs thereto and optional inputs and controls.

[0035]FIGS. 9a and b depicts one embodiment of an optional displayindicator for use in the present invention.

[0036]FIG. 10 depicts one embodiment of a liquid crystal displayindicator for use in the present invention.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

[0037] The accompanying Figures and Appendices A-C depict and describeembodiments of the apparatus or system of the present invention, andfeatures and components thereof, and set forth steps in the method ofthe present invention (see particularly Appendix C, with respect to oneexemplary embodiment). With regard to means for fastening, mounting,attaching or connecting the components of the present invention to formthe apparatus as a whole, unless specifically described otherwise, suchmeans are intended to encompass conventional fasteners such as machinescrews, machine threads, snap rings, hose clamps such as screw clampsand the like, rivets, nuts and bolts, toggles, pins and the like.Components may also be connected adhesively, by friction fitting, or bywelding or deformation, if appropriate. Electrical functions may beprovided by available or typical electrical components, including chipor board components. Connections may be soldered or potted, and wire,cable, wireless, optical fiber or other connections and junctions andelectrical components may be used. Unless specifically otherwisedisclosed or taught, materials for making components of the presentinvention are selected from appropriate materials such as metal,metallic alloys, natural or synthetic fibers, plastics and the like, andappropriate manufacturing or production methods including casting,extruding, molding and machining may be used.

[0038] Any references to front and back, right and left, top and bottom,upper and lower, and horizontal and vertical are intended forconvenience of description, not to limit the present invention or itscomponents to any one positional or spacial orientation.

[0039] Referring then to the Figures, FIG. 1 illustrates an airfoil 14as is typically used on air vehicles, including aircraft such asairplanes, jets and the like (not shown). The depicted airfoil isintended to airfoils for use on any vehicle having an airfoil orstructure which creates lift. In the instance of fixed wing aircraft, ahigh lift device 16 is typically attached to the airfoil (FIG. 1 shows aflap 16 movable to selected positions, including the position shown inphantom), and other high lift devices (not shown, but well known tothose skilled in the art) may also be attached to the airfoil 14 such asleading edge flaps. Such high lift or lift varying devices maysignificantly change the shape and aerodynamic characteristics of theairfoil 14. A proximity position sensor 18 is attached either directlyto the high lift device 16 or to the mechanism that moves the high liftdevice 16 to indicate its position. The airfoil 14 consists of aninterior compartment 20 at least part of which is generally free ofstructure.

[0040] In accordance with one embodiment of the present invention, oneor more orifices 22, 24 are located on and extend through the airfoilexterior surface 26 and provide communication of pressure or pressuresfrom the exterior to the compartment 20. The upper orifice 22communicates pressures from or adjacent to the upper airfoil surface 30and is located along the upper surface 30 of the airfoil 14 where thelocal pressure varies with angles from zero lift. More than one upperorifice may be provided, as indicated by phantom rear upper orifice 22′.In this instance the orifices may be operably coupled or bussedtogether. There will be many locations where the local pressures varyuniquely with angles from zero lift along the airfoil surface 30 and thepositional and dimensional relationships shown in FIG. 1 are intended tobe representative of any such location. For example, generally front andrear orifices 22 and 22′ could be used on the upper surface 30 of theairfoil 14.

[0041] The upper orifice 22 communicates to the interior of an air waterseparator 34. The interior of the air water separator 34 communicates tothe barbed hose adapter 36 which is secured to the side of the air waterseparator 34 and sealed with a typical washer or like sealing means. Thebarbed hose adapter 36 communicates to a flexible tube or hose 38. Theflexible hose 38 communicates to a suitable pressure sensor or the likelocated in or adjacent to the angle from zero lift computer (not shownliterally, but see FIG. 4 for a representation thereof). A quick drain40 is fastened to the bottom of the air water separator 34 and isaccessible through a hole or grommet 42 in the bottom surface 46 of theairfoil 14. The quick drain 40 allows water to be drained from the airwater separator 34 by overcoming spring pressure by pressing on theflange or grommet 42 in an upward direction and is normally sealed whenthe spring pressure is not overcome. A threaded plug (not shown) wouldalso serve a similar purpose. The air water separator 34 may be selectedand/or constructed depending upon the wing construction type.

[0042] Referring to FIG. 3, for composite wing structures, an upper airwater separator tube 48 slides within the lower tube 50 of the air waterseparator allowing the length of the air water separator to be adjustedto the thickness of the airfoil. Structural adhesive may be used to bondand seal the structural ring 52 to the inside of the upper airfoilsurface, to bond and seal the upper tube to the ring and bond and sealthe lower tube 50 to the upper tube 48.

[0043] Typical installation of the upper air water separator tube 48 maybe outlined as follows. The air/water separator is installed under theupper wing pressure port and is designed to prevent water or debris thatenters the upper wing port from getting into the tubing leading to atransducer or microprocessor (as explained below). The air/waterseparator is drainable through the bottom skin of the wing. The purposeof the telescoping air/water separator shown is to simplifyinstallations, especially retrofits, and to prevent having to bore a 1″OD installation hole in the bottom skin of the wing, but other air/waterseparator or drains may be used as well. First, gain access to the innerwing where the air/water separator is to be installed. This may be doneby removing the wing tip, or through an access panel. Sometimes accesscan be gained through the nav light lens. Plan the installation so thatthe air/water separator is normal (perpendicular) to the bottom skin ofthe wing and meets the upper wing skin at the targeted % chord. Thebrass barb should be oriented near the top wing skin. It is recommendedthat the inner wing skin bid and core be removed if possible subject tostructural considerations. Due to variables with the manufacture of thephenolic tubes, it may be necessary to sand the outer diameter of thesmaller inner phenolic tube until it easily slides within the largerphenolic tube. This can be done by hand or on a grinder. Insert theinner tube within the larger phenolic tube. Bevel the smaller diametertube so that it is flush with the upper wing skin and the brass barb isoriented the proper direction for easy removal and installation. Thelarger tube should not be beveled because it is perpendicular to thebottom wing skin. The tubes should be trimmed to the proper length.Allow enough length in the larger diameter phenolic tube for thealuminum donut to slide far enough into it without contacting thesmaller phenolic tube for a flush installation. It may be a good idea toorient the hose barb so that it can be removed or replaced. Test fit theinstallation by elongating the telescoping air/water separator into thearea where the core was removed. If the fit is good, use epoxy flox orthe equivalent and bond into position and seal the area where the twotubes meet. Drill through the bottom skin of the wing into the center ofthe phenolic tube. Enlarge the opening to the inner diameter of theair/water separator. Drill through the upper skin using a #60 drill bit(0.040″) into the center of the air/water separator. The hole should bedrilled normal (perpendicular) to the wing skin. Remove any burs causedby the drilling process. The upper wing skin in the area of the holeshould not be excessively thick (probably not more than ⅛″). Ifnecessary, accessing through the bottom of the air/water separator, usea ¼″ or larger long drill bit to remove any excess thickness in theupper wing skin just under the pressure port, being careful not to drillall the way through the upper wing skin. Install the drain assembly intothe AL donut using Teflon tape or equivalent as a thread sealer. If itwill not screw into the donut from one direction try the other. Sand theouter diameter of the aluminum donut until it snugly fits within thephenolic air/water separator. Do a test fit insuring the assembly willslide far enough up into the air/water separator so that the drain isflush with the bottom wing skin. Use epoxy or structural adhesive tobond the AL donut into the air/water separator. The push to drainassembly should be flush with the bottom skin of the wing. Pressure testthe entire assembly using soapy water and fix any leaks. Referring toFIG. 7, for aluminum or fabric wing structures 56 an upper air waterseparator block 58 is bonded to the interior of the upper wing skin andconnects to a flexible tube 60 which connects to a lower block 62attached to the lower interior wing skin.

[0044] Referring to FIG. 2, the lower orifice 24, which communicatespressures from the lower surface or underside 66 of an airfoil 14, islocated along the lower surface 66 of the airfoil where the localpressures vary uniquely with angle from zero lift. As in the instance ofthe upper surface 30 of an airfoil, there will be many locations wherethe local pressure varies uniquely with angle from zero lift and thepositions and dimensions depicted in FIG. 2 are not intended to beexclusive. The lower orifice 24 communicates to a barbed hose adapter68, in turn coupled to or communicating with a flexible hose 70. Theflexible hose 70 communicates to a pressure sensor located in the anglefrom zero lift computer (not shown, but see FIG. 4 for arepresentation). Structural adhesive is used to seal and bond thestructural ring 72 to the inside of the lower airfoil surface. The barb68 is fastened to the ring 72.

[0045] Pressures from the pitot and static system (ports, probes or thelike not shown, but typical and well-known to those skilled in the art)may be communicated to the angle from zero lift computer via pressuretubes of suitable size that are tied into the air vehicle pitot andstatic system or from the air vehicle pitot tube and static port. Anysuitable method to accomplish this would be acceptable.

[0046] The airfoil pressures from the upper and/or lower orifice ororifices are connected to a pressure sensor or sensors 80 located in oradjacent to the angle from zero lift computer 82 as illustrated in FIGS.5 and 7. The pitot and static pressures are also connected to a pressuresensor or sensors 80 as illustrated in FIGS. 5 and 7. Referring to FIG.5, the pressure sensors 80 output a voltage which is communicated to ananalog to digital converter 84. The analog to digital converter 84communicates digital data electronically to a microprocessor 86. Itshould be noted that the pressure sensor(s) 80 and the analog to digitalconverter 84 may be located immediately adjacent to or remote from theangle from zero lift computer 82.

[0047] The high lift device proximity sensor 18 communicateselectrically the position of the high lift device 16 to a microprocessor86 located within the angle from zero lift computer 82. Three momentarypush switches 88, 90, 92 communicate with the microprocessorelectrically. The mode switch 88 is used to select the calibratefunction of the angle from zero lift system and is used to select one ofseveral modes as will be explained herein below. The record switch 90 isused to record data in the EEPROM which stores data and is non-volatile.The test switch 92 is used to put the angle from zero lift system into atest mode.

[0048] An electrical port 94, which may be described as an opencollector sink, is connected to a power transistor 96 which is connectedto the microprocessor 86 and may be used as a current sink to activateexternal electrical devices when the airfoil's angle from zero lift isapproaching an unsafe angle from zero lift where the airflow willseparate from the airfoil.

[0049] A voice playback system, including, for example, chip 100 may beconnected to the microprocessor 86 and may be amplified 102 andavailable on the output pins of the angle from zero lift computer 82 ina high and low impedance conforming to air vehicle impedance standards.The voice playback system enunciates verbally when the angular deviationfrom angle of zero lift is approaching an unsafe angle from zero liftwhere the airflow will separate from the airfoil, and enunciatesverbally when system errors have been detected.

[0050] A tone playback system may be connected to the microprocessor,and may be amplified and available on the output pins of the angle fromzero lift computer 82. The tone system plays a middle “C” tone as areference tone during approaches to landing. Another tone is played overthe middle “C” tone and is bent from a middle “C” in either directiondepending upon the relative angle from zero lift at the optimum anglefrom zero lift for approaches to landing. The reference and bent middle“C” tones used in this application are not limiting as other tones otherthan “C” would work as well.

[0051] An electrical communications port 106 may be available on theoutput pins of the angle from zero lift computer 82. This port 106 maybe used to export angular deviation from angle of zero lift related datato other devices having compatible communications ports. Othercommunication standards are available and the exemplary port 106 (anRS-232 port) used in this embodiment is not limiting.

[0052] As shown in FIG. 5, a EEPROM 110 is a part of the angle from zerolift computer 82. The EEPROM 110 electrically communicates with themicroprocessor 86 and is used to permanently store digital data relatingto the flight calibration processes.

[0053] The angle from zero lift indicator 112 depicted in FIGS. 6 and 10preferably is a dash or console mounted liquid crystal display device112 and, as shown in FIG. 8, electrically communicates with themicroprocessor 86. Note that FIG. 8 also depicts optional inputs,outputs and controls, such as flap switches 113, landing gear switch 115and grounds 117. The indicator 112 has a digital numeric display 114 atthe bottom of the display controlled by the microprocessor 86 andconsisting of three digits with seven segments each that indicatesangles from zero lift in units of angular deviation from angle of zerolift. The indicator has 27 colored bars 116 each individually controlledby the microprocessor 86 that are either turned on or off and provide apilot with an analog type display of the angular deviation from angle ofzero lift relative to zero lift. The indicator has a green donut shapedsegment 118 in the center of the display that when turned on indicatesthat the airfoil 14 is configured for an approach to landing. Thedisplay 112 has a green split segment 120 several bars below the donut118 that indicates the angular deviation from angle of zero lift wherethe airfoil is at a performance angular deviation from angle of zerolift, for example, at the best lift over drag angular deviation fromangle of zero lift or the maximum endurance angular deviation from angleof zero lift. The chevron shaped segments 122 at the top of theindicator are red and indicate an angular deviation from angle of zerolift that is close to where the air may separate from the airfoil. Averbal warning message enunciates continually when the indicatordisplays an angular deviation from angle of zero lift in the redchevrons (FIG. 6a). The optimum angular deviation from angle of zerolift for an approach to landing is displayed when the yellow segmentsand red chevrons are turned on abeam and above the donut (FIG. 6b). Theoptimum performance angular deviation from angle of zero lift isdisplayed when the green segments, yellow segments, and red chevrons atand above the split green segment are turned on (FIG. 6c). The whitesegment under the digital display, all the green segments, all theyellow segments and all the red segments are turned on when the airfoilis at zero degrees from zero lift (FIG. 6d). The colors, number, andspecific arrangement of the segments on the display described above arepreferred, but others may be used as well.

[0054] An alternative display 140 as shown in FIGS. 9a and b compriseseight colored light emitting diodes. The top four diodes 142 are red incolor, the middle three diodes 144 are amber in color and the bottomdiode 146 is green. The light emitting diodes are arranged in a row andmay be called a LED ladder. FIG. 9b depicts a typical assembly wherein abezel 143 houses the diodes which are held in place by a suitable edgeconnector 145. At high angles from zero lift where the airfoil is closeto the critical angle one or more of the red LED's and all the otherLED's are illuminated. At the optimum angle from zero lift for anapproach the middle amber colored LED and all LED's below areilluminated. As the angle from zero lift increases from zero angle theLED ladder illuminates from the bottom green up the ladder to the redLED's.

[0055] Either of the two display embodiments and other alternativeembodiments may be dimmable, in part or as a whole, by using a dimmerpotentiometer (not shown) which controls a voltage regulator (not shown)to adjust the current through, for example, the display backlight bulbs.Preferably, a wide range of brightness will be available. Dimming of theeight bulb ladder LED display 142 would be accomplished, for example, bypushing a dimmer button or switch (not shown) which may either affectthe display continuously giving a selectable level of brightness, ordrop it to a preselected night viewing level of brightness. See FIG. 8for coupling of the optional dimmer 119.

[0056] The apparatus of the present invention may be adapted to assessand process wing-tip pressure measurements. No matter where they are,one or more of the pressure ports or pressure sensing openings, e.g.,ports 22, 24, may be adapted to include an electronic device at theorifice, e.g., a pressure transducer, an analog to digital converterand/or conditioner (not shown, but well-known to those skilled in theart). In this instance, the digital equivalent of sensed pressure wouldbe communicated to the processor 82. The present invention may beadapted to provide additional information to a pilot, such as a landinggear warning signal (see FIG. 8). The present invention may be built-inin new aircraft or it may be retrofit in existing aircraft. Wing and/orpressure ports or openings and optional air/water separators, ifrequired, may be installed through the wing tip or in areas accessiblethrough existing access panels. The ports are typically andapproximately the diameter of a common paper clip and are located at 10%chord or aft. For the routing of pressure tubes from the wing pressureports to the cabin area, appropriate bulkhead and other fittings and/orjunctions may be used. The CPU 82 selected for use with the presentinvention may be located behind the instrument panel or anywhere in theaircraft or cabin. The display instrument may be panel or glare shieldmounted. Appendix B is provided to give an example of components, andspecifications thereof, for a typical commercially available embodimentof the apparatus of the present invention, and Appendix A provides themanufacturer and model number of some of the components.

OPERATION

[0057] While the operation of the angle from zero lift measuring anddisplaying system will be readily apparent to those skilled in the art,the following discussion of the operation of one embodiment of thepresent invention will be presented for convenience and understanding.

[0058] Referring FIG. 4, the pressure at two orifices on the airfoil,block 150, are communicated to a temperature compensated differentialpressure transducer, block 152. Two additional pressures from the pitotand static systems, block 154, installed on all air vehicles are alsocommunicated to a temperature compensated differential pressuretransducer 156.

[0059] The pressure transducer electrical outputs are conditioned anddigitized and communicated to a processor 158 where the below describedcalculations are performed. The differential pressure transducer has aunique error called the zero pressure offset, it does not necessarilyoutput exactly 0 volts at zero p.s.i. The digital differential pressurefrom each pressure transducer is adjusted to compensate for each zeropressure offset using data recorded in the EEPROM by subtracting theEEPROM data from the saved pressure data.

P _(ADJ) =P _(X) −P _(X) HC

[0060] The data in the EEPROM resulted from taking a reading in thesystem's hangar calibration mode while the airfoil was in a hangar orthe like where the air is not in motion relative to the pressure ports.

[0061] The aircraft (i.e., the airfoil) is then flown into the air and azero G maneuver is accomplished (or this may be simulated) whichindicates where no lift, either positive or negative, is being generatedby the aircraft. The systems computer is put into the zero lift modeusing the mode switch and the sensed pressures are saved in themicroprocessor 82 at the appropriate time using the record switch. Inthe case of using differential pressure transducers, the pressures fromthe differential pressure transducers are adjusted for the zero offsetsand saved in the microprocessor 82. A coefficient of pressure iscalculated using all four pressures where:

CP=(P _(W1) −P _(W2))/(P _(P) −P _(S))

[0062] The coefficient of pressure is communicated from themicroprocessor to the EEPROM. The EEPROM saves the coefficient ofpressure in non-volatile memory.

[0063] The airfoil is then flown into the air (or this situation issimulated) and a angle from zero lift is flown using the airspeedindicator as a reference where it is desirable that the angle advisorysystems activate. The system computer is put into the angle advisorymode using the mode switch and the pressures saved at the appropriatetime using the record switch. The pressures from the differentialpressure transducers are adjusted for the zero offsets and saved in themicroprocessor. A coefficient of pressure is calculated using all fourpressures where:

CP=(P _(W1) −P _(W2))/(P _(P) −P _(S))

[0064] The coefficient of pressure is communicated from themicroprocessor to the EEPROM. The EEPROM saves the coefficient ofpressure in non-volatile memory.

[0065] As indicated generally at block 160, the above-outlinedoperations provide two coefficient of pressure flight data points which,along with the hanger calibration, are sufficient for each airfoilconfiguration to build the relationship between the coefficient ofpressure and angle from zero lift. It is well known by those skilled inthe art that the coefficient of pressures vary uniquely with angles fromzero lift and that as little as two data points could define all thepoints over the entire coefficient of pressure and angle from zero liftrelationship.

[0066] Preferably, the angle from zero lift instrument display 112displays not only the full range of angles from zero lift, as describedherein above the display 112 also has special markings to indicate theoptimum approach to landing angle from zero lift (donut) and the optimumperformance angular deviation from angle of zero lift (split light bar)for each selected airfoil configuration. The display 112 is designed sothat these special markings are located on the indicator so that theirrelative position to the angle advisory and angle from zero lift lightbars are typical of most airfoils.

[0067] Should the approach angular deviation from angle of zero lift berequired and the approach angle from zero lift is not in preciseagreement with the display, another data point may be recorded to theEEPROM via calculations explained below, will force the display intoexact agreement by flying at the desired angle for approach using theairspeed indicator as a reference. The systems computer is put into theapproach advisory mode using the mode switch and the pressures recordedat the appropriate time using the record switch. The pressures from thedifferential pressure transducers are adjusted for the zero offsets andsaved. A coefficient of pressure is calculated using all four pressureswhere:

CP=(P _(W1) −P _(W2))/(P _(P) −P _(S))

[0068] The coefficient of pressure is communicated from themicroprocessor to the EEPROM. The EEPROM saves the coefficient ofpressure in non-volatile memory.

[0069] Should the performance angular deviation from angle of zero liftbe required and the performance angle from zero lift is not in preciseagreement with the display, another data point may be recorded to theEEPROM, which via calculations (explained herein below) will force thedisplay into exact agreement by flying at the desired angle forperformance using the airspeed indicator as a reference. The systemscomputer is put into the performance advisory mode using the mode switchand the pressures recorded at the appropriate time using the recordswitch. The pressures from the differential pressure transducers areadjusted for the zero offsets and saved. A coefficient of pressure iscalculated using all four pressures where:

CP=(P _(W1) −P _(W2))/(P _(P) −P _(S))

[0070] The coefficient of pressure is communicated from themicroprocessor to the EEPROM. The EEPROM saves the coefficient ofpressure in non-volatile memory.

[0071] Having saved the hangar calibration zero offset in the EEPROM,the zero degree angle from zero lift in the EEPROM, and the advisoryangle from zero lift in the EEPROM, there is enough data points to fullydescribe the relationship between the coefficient of pressures and theangles from zero lift. There may be two additional optional data pointsstored in the EEPROM for the current airfoil configuration underconsideration. The microprocessor sensing a specific airfoilconfiguration uses the appropriate data points stored in the EEPROM andcomputes AOA using acceptable mathematical techniques to connect thedata and extending beyond the data points.

[0072] The flight mode is the normal operating mode for the system. Inthe flight mode, the airfoil is then flown at any angular deviation fromangle of zero lift into the air. The pressures from the differentialpressure transducers are adjusted for the zero offsets and saved in themicroprocessor 82. A coefficient of pressure is calculated using allfour pressures where:

CP=(P _(W1) −P _(W2))/(P _(P) −P _(S))

[0073] The coefficient of pressure is used to calculate the angulardeviation from angle of zero lift, block 161. At block 162, the angulardeviation from angle of zero lift is displayed on the angle from zerolift instrument and communicated to the data port. Should the angulardeviation from angle of zero lift exceed the advisory angular deviationfrom angle of zero lift, an aural warning sounds repeatedly and the opencollector sinks to ground. The system while in the flight modecontinually loops through the aforementioned flight mode processcontinually updating the angle from zero lift display and the data port.

[0074] Preferred calibrating of the instrument involves recording theresults for two airfoil configurations, one configuration for the highspeed cruise configuration and one configuration for the approach tolanding configuration. On some airfoils, there is the option to selectmany different configurations in-between the cruise and the approach tolanding configuration. In this event, there exists an option for thosein-between configurations to assume that one known configuration willproduce a more conservative display than the other configuration. Thisassumption is set either in software directly or by physically settingan assumption jumper switch located on the systems computer. For thosein-between configurations, the assumption jumper will communicate to themicroprocessor the assumption desired, either to use the cruise or theapproach to landing EEPROM data to compute the AOA for the in-betweenairfoil configuration selected. The system does however allow more thantwo airfoil configurations if opted with the ability to estimate theangular deviation from angle of zero lift using known coefficients ofpressures for a multitude of airfoil configurations and the saved EEPROMdata points for two or more airfoil configurations.

[0075] For additional information and details regarding the operationand software program flow of the present invention, the reader isreferred to Appendix C, outlining in block summary form the program flowof one embodiment of the invention.

[0076] The system and method of the present invention may be adapted toinclude an automatic self-test upon power up. Also, if desired, theinstrument may be adapted for testing at any time by providing a testsequence activating button or switch and appropriate hardware andsoftware.

[0077] The present invention may be embodied in other specific forms andsteps without departing from the essential spirit or attributes thereof.It is desired that the embodiments described herein be considered in allrespects as illustrative, not restrictive, and that reference be made tothe appended claims for determining the scope of the invention.

[0078] For example, as set forth herein above, in one embodiment, thepresent invention relates to a measurement, data manipulation andinformation display system and method for use on air vehicles andcomprises a method of determining information for an air vehiclecomprising the steps of providing a first coefficient of pressure datapoint, providing a situation involving the air vehicle, derivingpressures during the situation and deriving a second coefficient ofpressure data point from the pressures derived during the situation, anddefining a line passing through the first and second coefficient ofpressure data points, whereby a selected coefficient of pressure datapoint on the line corresponds to the information. As set forth hereinabove, the data points may comprise coefficient of pressure data, andthe information is related to aircraft performance or flightinformation, for example, information relating to angles from zero lift.Thus, in one embodiment, the present invention comprises a method ofdetermining information for an aircraft comprising the steps ofproviding first input data, providing second input data, and processingthe first input data and said second input data according to apredetermined set of instructions to determine a third data point,wherein the third data point correlates to the information. In someembodiments, providing the second input data comprises flying theaircraft, and in some embodiments, providing the first input data isselected from the group comprising flying the aircraft, simulating theflying of the aircraft or assuming the first input data. The informationmay comprise aircraft information, such as flight information,including, but not limited to flight characteristics, performance, angleof attack, lift, yaw and the like. The method, including the processing,may be carried out by a microprocessor.

[0079] In another illustrative embodiment, the system and method of thepresent invention relates to a measurement, data manipulation andinformation display system and method for use on air vehicles andcomprises or may involve information relating to angles from zero yaw.(As used herein, the term “yaw” is intended to have its common meaning,as concerning aircraft, airplanes, spacecraft or projectiles, i.e., toturn by angular motion about the vertical axis.) In this embodiment, thesystem invented involves determining information for an air vehiclecomprising the steps of providing a first coefficient of pressure datapoint, providing a situation involving the air vehicle, derivingpressures during the situation and deriving a second coefficient ofpressure data point from the pressures derived during the situation, anddefining a line passing through the first and second coefficient ofpressure data points, whereby a selected coefficient of pressure datapoint on the line corresponds to the information. In this embodiment,the information may relate to angles from zero yaw, and the method maycomprise the steps of creating a zero yaw situation, deriving pressuresfrom the zero yaw situation, and deriving a first coefficient ofpressure data point from the pressures derived during the zero yawsituation, creating a yaw producing situation, deriving pressures duringthe yaw producing situation and deriving a second coefficient ofpressure data point from the pressures derived during the yaw producingsituation, and defining a line passing through the first and secondcoefficient of pressure data points, whereby a selected coefficient ofpressure data point on the line corresponds to an angle of attack. Inthis embodiment the method further comprises a sensor or sensorsappropriately located on an air foil of an air vehicle. It should beappreciated that the accompanying Figures depict and diagram apparatusand method for use in measurement, data manipulation and informationdisplay systems and methods for use on air vehicles, for example, forlift and/or yaw. Thus, what is referred to as the “lift computer”(represented in FIG. 4) may also be referred to as simply the “computer”or “processor” or the “lift and/or yaw computer.” Further, it may beused to process, manipulate, record and output aircraft informationother than lift or yaw information.

[0080] It should be appreciated that one embodiment of the presentinvention, whether concerning lift, yaw or another type of aircraftinformation, comprises a method of determining information for anaircraft comprising the steps of providing first input data, providingsecond input data, and processing the first input data and said secondinput data according to a predetermined set of instructions to determinea third data point, wherein the third data point correlates to theinformation. In some embodiments, providing the second input datacomprises flying the aircraft, and in some embodiments, providing thefirst input data is selected from the group comprising flying theaircraft, simulating the flying of the aircraft or assuming the firstinput data. The information may comprise aircraft information, such asflight information, including, but not limited to flightcharacteristics, performance, angle of attack, lift, yaw and the like.The method, including the processing, may be carried out by amicroprocessor.

[0081] In one embodiment, the apparatus of the present invention isprovided for measuring and displaying angular deviation from an angle ofzero lift or yaw for air vehicles with an air foil and comprises an airfoil pressure sensor on a surface of an air foil for sensing a local airfoil pressure, a micro processor operably coupled to the air foilpressure sensor for receiving and processing sensed pressures, forderiving a coefficient of pressure, and for relating said coefficient ofpressure to an angular deviation from angle of zero lift or yaw. Theapparatus may further comprise an electrical indicator operably coupledto the micro processor for communicating information about the angulardeviation. It may further comprise coupling the microprocessor to apitot and/or static pressure sensors associated with or carried by theair vehicle.

[0082] As set forth herein above, the present invention may be used tocreate and develop coefficient of pressure data points based onsimulation situations, or by using computational air foil data analysis.In other words, the zero lift or yaw situation may be an actualsituation involving the aircraft, a situation involving the aircraftwhich simulates an actual situation (e.g., a wind tunnel situation) orit may be an assumed, hypothesized or otherwise created situation. Thus,for some types of aircraft, the zero G maneuver may not be performed andthe coefficient of lift and/or yaw may be assumed to be zero, making thecalibration process even easier. At least one other data point or moreat any higher angle from zero lift and/or yaw is obtained, and therelationship between the coefficient of pressure and angle from zerolift or yaw is defined for a specific aircraft configuration. The anglefrom zero lift or yaw system instrument's processor uses variousacceptable mathematical functions to connect the points recorded innon-volatile memory providing the angular deviation from angle of zerolift or yaw for any coefficient of pressure resulting from anycombination of pressures at the pressure ports for a known aircraftand/or airfoil configuration. For in-between configurations, one or moresuitable methods of interpolation may be used to compute angle from zerolift or yaw. Thus, the method and apparatus of the present invention maytake into account, be adapted to, or otherwise be compatible withvarious aircraft and/or airfoils in which the configuration of theairfoil may be varied, for example, by using high lift devices. For aselected airfoil or aircraft configuration, a second, third or n sets ofcalibration data may be used.

[0083] A problem addressed by the present invention, in addition tothose mentioned herein above, is that probes of the type mentioned inthe above “Background” section must be located well ahead of the wing toreduce the effect of changes in lift and configuration. Usually theprobe is located on or near the nose of an aircraft. This is a problemsince many factors can affect the relationship between the local AOA andthe true wing AOA. The airflow around the nose is not the same as thatat the wing. Pitch rate errors occur when the nose of an aircraft ispitched up or down causing the probe to indicate too high or too low. Ina turn, for example, the nose is pitched up reducing the local flowangle and causing the AOA reading to be too low. Because of the shape ofthe nose, sensitivity errors are introduced. A one degree change in truewing AOA may cause the local flow angle at the nose to be 1.5 to 2degrees. In some embodiments, the present invention reduces oreliminates pitching, sensitivity and other errors common to traditionalprobe-based sensing devices.

[0084] Another problem with systems such as those noted above is thatthey do not account for varying aircraft configurations such as winghigh lift devices (e.g., flaps and slats) being extended or retracted.The present invention does and, for various aircraft configuration, asecond or third set of calibration data points may be provided. Once thedata points are stored in the system memory, all other coefficient ofpressures can be equated to a specific angular deviation from angle ofzero lift or yaw for a specific, selected aircraft configuration. Thus,in one embodiment, the present invention provides an apparatus andmethod for use on aircraft and includes pressure sensors, an aircraftconfiguration sensor, a processor, at least one data port forelectrically exporting angles from zero lift and/or yaw, and anindicator for visibly displaying angles from zero lift and or yaw.

[0085] In some embodiments, the present invention may sense, processand/or provide information relating to an airfoil approaching an unsafeangle from zero lift or yaw at which the airflow may separate from theairfoil or at which structural damage may occur.

[0086] In one embodiment, the present invention relates to a measurementand display system for use on air vehicles having an indicator visiblydisplaying angles from zero yaw and having a data port for electricallyexporting angles from zero yaw. In one embodiment, pressures from upperand lower wing ports and pressures from typical pitot in static ports ofair vehicle are converted to digital data and mathematically dividedinto one another. The result is a coefficient of pressure (CP) whichvaries uniquely with angles from zero yaw of the air vehicle. Theangular deviation from angle of zero yaw display is designed for airvehicles and is located in the cockpit and displays angles from zero yawon a digital display. The data port electrically exports angles fromzero yaw to other electronic devices using a compatible communicationsport. It has been discovered that if the coefficient of pressure datapoint for zero degrees angular deviation from angle of zero yaw and aminimum of one other coefficient of pressure data point is storedpermanently in the system's non-volatile memory, other angles from zeroyaw can be determined given any coefficient of pressure. For any givenairfoil, it has been discovered that a minimum of two or morecoefficient of pressures will accurately define the relationship betweenthe coefficient of pressure and the angular deviation from angle of zeroyaw. Once the data points are stored in the system's non-volatilememory, all other coefficient of pressures can be equated to a specificangular deviation from angle of zero yaw. Thus, in one embodiment, thepresent invention provides an apparatus and method for measuringpressures, for deriving and storing in non-volatile memory two or morecoefficient of pressure data points, and for exporting and displayingthe angles from zero yaw.

[0087] In one embodiment of the present invention, a microprocessor maybe used to permanently store the two or more coefficient of pressureswhich may be copied and installed in the angular deviation from angle ofzero lift and/or yaw system of similar air vehicles of the same type toaccurately display angles from zero lift and/or yaw.

[0088] It should be appreciated that the orifices 22, 24 (etc.) of thepresent invention may provide additional aircraft or airfoilinformation. For example, an abnormally low pressure reading or afalling pressure trend may indicate that an airfoil surface is becomingcontaminated, e.g., by icing.

[0089] In one embodiment, particularly as to the display 140 depicted inFIGS. 9a and 9 b, the diodes may comprise three red diodes, three amberdiodes and two green diodes.

[0090] With respect to the operation of the present invention, it is asis described herein above, and whenever “lift” is referred to in thedescription, it should be appreciated that yaw may be referred toalternatively or conjunctively as well. That is, the present inventionmay be used to sense, assess, determine and/or predict aircraftinformation, including information related to flight characteristicssuch as lift or yaw, and it may do so selectively as to either lift oryaw or both. It should also be appreciated that, in some embodiments,coefficients of pressure may be calculated using three pressures(instead of four) where

CP=(P _(W) −P _(S))/(P _(P) −P _(S))

[0091] where P_(W) is either P_(W1) or P_(W2)

[0092] Again, the present invention may be embodied in other specificforms and steps without departing from the essential spirit orattributes thereof. It is desired that the embodiments described hereinbe considered in all respects as illustrative, not restrictive, and thatreference be made to the appended claims for determining the scope ofthe invention. APPENDIX A Description Manufacturer Part # 5 psidifferential pressure sensor SenSym SDX05D4-A RS232 transmitter/receiverMaxim MAX232ACPE 8 bit microprocessor Microchip PIC16C65-A-20I/P −40 to+85 deg EPROM 8 bit × 128 Microchip 93LC46A-I/P Triple 2 chan MuxNational Semi CD4053BCN crystal @ 8MH DigiKey CTX-082 A/D converterw/serial interface Newark TLC1549CP Voice playback chip NuHorizons ISDISD1416PI 4 color liquid crystal display Crystalloid AOA LCD 32 segmentLCD driver Microchip AYO438/L steel drain valve Stratoflex 1/8NPT drainvalve

SPECIFICATIONS

[0093] Computer (CPU):

[0094] 12 to 28 volts DC

[0095] 0.5 ampere current draw maximum including backlight for displayinstrument

[0096] weight: 13 ounces including tray or cage

[0097] Auxiliary open collector max sinking 0.8 amps at 12-28 VDC

[0098] Display instrument (LCD):

[0099] weight: 2 ounces

[0100] backlight may be dimmable

[0101] four colors display

[0102] Pressure Sensors:

[0103] 5 p.s.i. operating

[0104] 20 p.s.i. burst

[0105] Voice Output:

[0106] 600 ohm audio output

[0107] 26 ohm audio output

[0108] adjustable volume control APPENDIX B One Typical CommerciallyAvailable Embodiment of the Present Invention AOA CPU angle-of-attackcentral processing unit is en- closed in a metal tray AOA LCDangle-of-attack liquid crystal display is a four color display andbacklit in a custom black anodized bezel Dimmer with knob dimmer toadjust the backlight of the display Wiring Harness customized to yourspecifications Pressure tubing color coded ⅛ inch OD tubing red, blue,green and clear Air/Water Separator either glass kit or aluminum kitsupplied if needed AOA Pressure Port decal decal or stencil Push to Testswitch momentary switch which when pushed starts the self test “AOAPASS” Microswitch flap position sensing Nylon tees and adapters two teeswith adapters to ⅛” OD tubing for the pilot and static taps Hose barbs &bulkhead enough to complete installation fittings SS recessed cap screwsfor the LCD display panel mount Manual illustrated operating,installation and education manual in 3 ring binder

What is claimed is:
 1. A method for determining information for an airvehicle comprising the steps of: providing a first coefficient ofpressure data point; providing a situation involving the air vehicle,deriving pressures during the situation and deriving a secondcoefficient of pressure data point from the pressures derived during thesituation; and defining a line passing through the first and secondcoefficient of pressure data points, whereby a selected coefficient ofpressure data point on the line corresponds to the information.
 2. Themethod according to claim 1 , wherein the information comprises flightinformation.
 3. The method according to claim 1 , wherein providing thefirst coefficient of pressure data point comprises assuming the firstcoefficient of pressure data point.
 4. The method according to claim 2 ,wherein the flight information comprises angle from zero lift.
 5. Themethod according to claim 2 , wherein the flight information comprisesangles from zero yaw.
 6. The method according to claim 2 , wherein theflight information relates to a selected air vehicle configuration. 7.The method according to claim 1 , further comprising an additional datapoint, said data points defining a curvilinear relation betweencoefficients of pressure and the information.
 8. A method of determiningangle of attack information for an air vehicle comprising the steps of:creating a zero yaw situation, deriving pressures during the zero yawsituation, and deriving a first coefficient of pressure data point fromthe pressures derived during the zero yaw situation; creating a yawproducing situation, deriving pressures during the yaw producingsituation and deriving a second coefficient of pressure data point fromthe pressures derived during the yaw producing situation; and defining aline passing through the first and second coefficient of pressure datapoints, whereby a selected coefficient of pressure data point on theline corresponds to an angle of attack.
 9. The method according to claim8 , further comprising pressure sensors including an airfoil pressuresensor operably carried by an airfoil of the air vehicle, a pitotpressure sensor and a static pressure sensor.
 10. The method accordingto claim 9 , wherein deriving the first and second pressure data pointscomprises obtaining an airfoil differential pressure from the airfoilpressure sensor and dividing said airfoil differential pressure by apitot static differential pressure obtained from the pitot and staticpressure sensors.
 11. A method of determining angular deviations from azero yaw angle for an air vehicle including an airfoil comprising thesteps of: providing the air vehicle with pressure sensors operablylinked to a microprocessor; performing a zero yaw flight maneuver toprovide the zero yaw angle and sensing pressures during the zero yawflight maneuver; using the microprocessor to derive a coefficient ofpressure from the sensed pressures and recording said coefficient ofpressure in the microprocessor; acquiring a coefficient of pressure datapoint while the air vehicle is producing yaw, communicating thecoefficient of pressure data point to the microprocessor to define aunique relationship between the coefficient of pressure; and providingan angular deviation from the zero yaw angle for any coefficient ofpressure.
 12. The method according to claim 11 , further comprisingproviding angular deviation information to a pilot of the air vehicle.13. The method according to claim 12 , further comprising providingwarnings to the pilot when an angle from zero yaw is approaching anangle at which airflow separates from the airfoil.
 14. An apparatus formeasuring and displaying aircraft information for aircraft with anairfoil and with a pitot pressure sensor and a static pressure sensorcomprising: an airfoil pressure sensor on a surface of the airfoil forsensing a local airfoil pressure; a microprocessor operably coupled tothe pitot and static pressure sensors and to said airfoil pressuresensor for receiving and processing sensed pressures, for recording afirst coefficient of pressure and a second coefficient of pressure,wherein the microprocessor processes and correlates said first andsecond coefficients of pressure to provide the aircraft information; andan electrical indicator operably coupled to the microprocessor forcommunicating the information.
 15. A method of determining informationfor an aircraft comprising the steps of: providing first input data;providing second input data; and processing said first input data andsaid second input data according to a predetermined set of instructionsto determine a third data point.
 16. The method according to claim 15 ,wherein the third data point correlates to the information.
 17. Themethod according to claim 15 , wherein providing said second input datacomprises flying the aircraft.
 18. The method according to claim 15 ,wherein said providing said first input data is selected from the groupcomprising flying the aircraft, simulating the flying of the aircraft orassuming the first input data.
 19. The method according to claim 15 ,wherein the information comprises flight information.
 20. The methodaccording to claim 15 , wherein said processing is carried out by amicroprocessor.
 21. A method of calibrating an aircraft sensor forproviding aircraft information, comprising: providing a zero coefficientof pressure data point; and providing an aircraft situation, obtaining apressure reading during the aircraft situation and deriving a secondcoefficient of pressure data point from the pressure reading obtainedduring the aircraft situation.
 22. The method according to claim 21 ,further comprising using the zero coefficient of pressure data point andthe second coefficient of pressure data point to define a line, wherebya selected coefficient of pressure data point on the line corresponds tothe aircraft information.
 23. The method according to claim 22 , whereinsaid providing a zero coefficient of pressure data point comprisescreating a zero yaw situation and obtaining pressures during the zeroyaw situation.
 24. The method according to claim 22 , wherein saidproviding a zero coefficient of pressure data point comprises creating azero lift situation and obtaining pressures during the zero liftsituation.
 25. The method according to claim 21 , wherein an aircraftcarrying the aircraft sensor has a variable configuration.
 26. Themethod according to claim 21 , wherein an aircraft carrying the aircraftsensor has at least two configurations, and wherein the method may beperformed at two selected configurations, the method further comprisinginterpolating for configurations other than said two selectedconfigurations.
 27. The method according to claim 26 , wherein said twoselected configurations comprise flap settings.